Gearboxes for aircraft gas turbine engines

ABSTRACT

Gearboxes for aircraft gas turbine engines, in particular to arrangements for journal bearings such gearboxes, and to related methods of operating such gearboxes and gas turbine engines. Example embodiments include a gearbox for an aircraft gas turbine engine, the gearbox including: a sun gear; a plurality of planet gears surrounding and engaged with the sun gear; and a ring gear surrounding and engaged with the plurality of planet gears, each of the plurality of planet gears being rotatably mounted around a pin of a planet gear carrier with a journal bearing having an internal sliding surface on the planet gear and an external sliding surface on the pin.

The present disclosure relates to gearboxes for aircraft gas turbineengines, in particular to arrangements for journal bearings in suchgearboxes, and to related methods of operating such gearboxes and gasturbine engines.

Gas turbine engines with larger diameter fans may incorporate a gearboxconnecting the fan to a core shaft of the engine core. An advantage ofdoing so is that both the fan and the engine core can be designed tooperate efficiently as the fan size is scaled up, since the rotationalspeed of the fan is limited by the tangential speed of the fan tips. Thegearbox allows for a reduction in rotational speed of the fan comparedto that of the engine core, at the expense of additional weight of thegearbox and some efficiency losses within the gearbox. To maintainefficiency of operation of the engine, the gearbox needs to be designedto minimise weight and maximise efficiency. Bearings are a source oflosses within a gearbox, and therefore need to be optimised to seek tomaximise the efficiency of the gearbox.

According to a first aspect there is provided a gearbox for an aircraftgas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and a ring gear surrounding and engaged with the plurality        of planet gears, each of the plurality of planet gears being        rotatably mounted around a pin of a planet gear carrier with a        journal bearing having an internal sliding surface on the planet        gear and an external sliding surface on the pin,    -   wherein the internal or external sliding surface of the journal        bearing has a surface coating comprising a layer of an alloy        having aluminium or copper as a primary constituent.

The ring gear may have a pitch circle diameter of around 550 mm orgreater. Each of the planetary bearings may have a maximum operatingspecific load and a maximum operating sliding speed, wherein the maximumoperating specific load multiplied by the maximum operating slidingspeed is around 240 MPa m/s or greater. The maximum operating slidingspeed may be around 30 m/s or greater, and optionally no greater thanaround 60 m/s. The maximum operating specific load may be around 7 MPaor greater.

The maximum operating specific load multiplied by the maximum operatingsliding speed may be less than around 720 MPa m/s.

The surface coating may be provided on the external sliding surface ofeach journal bearing.

The external sliding surface of each journal bearing may be on a sleevemounted around a respective pin.

A thickness of the surface coating may be between around 40 and around200 micrometres.

A thickness of the layer may be between around 40 and around 100micrometres.

A gas turbine engine for an aircraft may comprise: an engine corecomprising a turbine, a compressor and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of blades; and a gearbox according to thefirst aspect, the gearbox configured to receive an input from the coreshaft and provide an output drive to the fan so as to drive the fan at alower rotational speed than the core shaft.

Where the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

According to a second aspect there is provided a method of operating thegas turbine engine, the method comprising operating the engine atmaximum take-off conditions, wherein for each journal bearing in thegearbox a specific loading multiplied by a sliding speed is greater thanaround 240 MPa m/s.

The specific loading multiplied by a sliding speed for each journalbearing may be less than around 720 MPa m/s.

According to a third aspect there is provided a gearbox for an aircraftgas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and    -   a ring gear surrounding and engaged with the plurality of planet        gears, each of the plurality of planet gears beings rotatably        mounted around a pin of a planet gear carrier with a journal        bearing having an internal sliding surface on the planet gear        and an external sliding surface on the pin,    -   wherein a ratio of a length, L, of the internal and external        sliding surfaces to a diameter, D, of the journal bearing is        between around 0.5 and 1.4.

The ring gear may have a pitch circle diameter of around 550 mm orgreater.

The L/D ratio in some examples may be between around 1.1 and 1.3.

Each of the planetary bearings may have a maximum operating specificload and a maximum operating sliding speed, wherein the maximumoperating specific load multiplied by the maximum operating slidingspeed is around 240 MPa m/s or greater.

The maximum operating specific load multiplied by the maximum operatingsliding speed may be less than around 720 MPa m/s.

The pitch circle diameter of the ring gear may be no greater than 1200mm.

A gas turbine engine for an aircraft may comprise: an engine corecomprising a turbine, a compressor and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of blades; and a gearbox according to thethird aspect, the gearbox configured to receive an input from the coreshaft and provide an output drive to the fan so as to drive the fan at alower rotational speed than the core shaft.

Where the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

According to a fourth aspect there is provided a method of operating thegas turbine engine, the method comprising operating the engine atmaximum take-off conditions, wherein for each journal bearing in thegearbox a specific loading multiplied by a sliding speed is greater thanaround 240 MPa m/s.

The specific loading multiplied by a sliding speed for each journalbearing may be less than around 720 MPa m/s.

According to a fifth aspect there is provided a method of operating agearbox for an aircraft gas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and    -   a ring gear surrounding and engaged with the plurality of planet        gears, the ring gear having a pitch circle diameter of around        550 mm or greater,    -   wherein each of the plurality of planet gears is rotatably        mounted around a pin of a planet gear carrier with a journal        bearing having an internal sliding surface on the planet gear        and an external sliding surface on the pin, an oil film between        the internal surface on the planet gear and the external sliding        surface on the pin varying between a maximum thickness and a        minimum thickness around the journal bearing,    -   the method comprising operating the aircraft gas turbine engine        at maximum take-off conditions such that the minimum thickness        of the oil film varies between the plurality of planet gears by        no more than 8% from a mean minimum oil film thickness.

A diameter, D, of each journal bearing may be between around 120 mm andaround 200 mm.

A length, L, of the internal and external sliding surfaces of eachjournal bearing may be between around 0.5 and around 1.4 of thediameter, D. The ratio L/D may be between around 1.1 and around 1.3.

The mean minimum oil film thickness at maximum take-off conditions maybe between around 3.5 and 8 micrometres.

An eccentricity ratio of each journal bearing during operation of thegas turbine engine at maximum take-off conditions may be within a rangeof between around 0.94 and 0.97.

For each journal bearing in the gearbox a specific loading multiplied bya sliding speed may be greater than around 240 MPa m/s.

The specific loading multiplied by a sliding speed for each journalbearing may be less than around 720 MPa m/s.

According to a sixth aspect there is provided a gearbox for an aircraftgas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and    -   a ring gear surrounding and engaged with the plurality of planet        gears,    -   each of the plurality of planet gears being rotatably mounted        around a pin of a planet gear carrier with a journal bearing        having an internal sliding surface on the planet gear and an        external sliding surface on the pin, an oil film between the        internal surface on the planet gear and the external sliding        surface on the pin varying between a maximum thickness and a        minimum thickness around the journal bearing,    -   and wherein, during operation of the aircraft gas turbine engine        at maximum take-off conditions, the minimum thickness of the oil        film varies between the plurality of planet gears by no more        than 8% from a mean minimum oil film thickness.

The various optional features mentioned above in relation to the fifthaspect may apply also to the sixth aspect.

A gas turbine engine for an aircraft may comprise: an engine corecomprising a turbine, a compressor and a core shaft connecting theturbine to the compressor; a fan located upstream of the engine core,the fan comprising a plurality of blades; and a gearbox according to thesixth aspect, the gearbox configured to receive an input from the coreshaft and provide an output drive to the fan so as to drive the fan at alower rotational speed than the core shaft.

Where the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

According to a seventh aspect there is provided a gearbox for anaircraft gas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and    -   a ring gear surrounding and engaged with the plurality of planet        gears, each of the plurality of planet gears being rotatably        mounted around a pin of a planet gear carrier with a journal        bearing having an internal sliding surface on the planet gear        and an external sliding surface on the pin,    -   wherein, during operation of the aircraft gas turbine engine at        maximum take-off conditions, a specific operating load        multiplied by an operating sliding speed of each journal bearing        is around 300 MPa m/s or greater.

The ring gear may have a pitch circle diameter of around 550 mm orgreater.

During operation of the aircraft gas turbine engine at maximum take-offconditions, the specific operating load multiplied by the operatingsliding speed of each journal bearing may be no greater than around 720MPa m/s.

During operation of the aircraft gas turbine engine at maximum take-offconditions, the sliding speed of each journal bearing may be greaterthan around 30 m/s or 35 m/s.

During operation of the aircraft gas turbine engine at maximum take-offconditions, the sliding speed of each journal bearing may be less thanaround 49 m/s, 47 m/s, 43 m/s or 40 m/s.

During operation of the aircraft gas turbine engine at maximum take-offconditions, the specific operating load of each journal bearing may bearound 5 MPa or greater.

During operation of the aircraft gas turbine engine at maximum take-offconditions, the specific operating load of each journal bearing may beless than around 20 MPa.

During operation of the aircraft gas turbine engine at maximum take-offconditions, the specific operating load of each journal bearing may begreater than around 10 MPa.

A diametral clearance of each journal bearing may be between around 1‰and around 2‰. The diametral clearance may be between around 1.4‰ andaround 1.6 ‰.

A gas turbine engine for an aircraft may comprise:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox according to the seventh aspect, the gearbox        configured to receive an input from the core shaft and provide        an output drive to the fan so as to drive the fan at a lower        rotational speed than the core shaft.

Where the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

According to an eighth aspect there is provided a method of operatingthe aircraft gas turbine engine, the method comprising operating theaircraft gas turbine engine at maximum take-off conditions such that aspecific operating load multiplied by an operating sliding speed of eachjournal bearing is around 300 MPa m/s or greater.

The various optional features relating to the seventh aspect may alsoapply to the eighth aspect.

According to a ninth aspect, there is provided a gearbox for an aircraftgas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and    -   a ring gear surrounding and engaged with the plurality of planet        gears, each of the plurality of planet gears being rotatably        mounted around a pin of a planet gear carrier with a journal        bearing having an internal sliding surface on the planet gear        and an external sliding surface on the pin,    -   wherein, with the aircraft gas turbine engine operating at        maximum take-off conditions, an eccentricity ratio of each        journal bearing, defined as 1-2H_(min)/c where H_(min) is a        minimum oil film thickness between the internal and external        sliding surfaces and c is the diametral clearance of the journal        bearing, is greater than around 0.84.

The ring gear may have a pitch circle diameter of around 550 mm orgreater.

With the aircraft gas turbine engine operating at maximum take-offconditions, the eccentricity ratio of each journal bearing may bebetween around 0.94 and 0.97.

The diametral clearance of each journal bearing may be between around 1‰and around 2‰. In some examples the diametral clearance of each journalbearing may be between around 1.4‰ and 1.6‰.

With the aircraft gas turbine engine operating at maximum take-offconditions, a temperature of oil flowing into each journal bearing maybe no greater than around 100° C.

With the aircraft gas turbine engine operating at maximum take-offconditions, a pressure of oil flowing into each journal bearing may bewithin a range from around 50 kPa to around 350 kPa.

An inefficiency of each journal bearing, defined as a percentage powerloss with the aircraft gas turbine engine operating at maximum take-offconditions, may be less than around 0.225%. In some examples theinefficiency of each journal bearing may be no less than around 0.1%.

A gas turbine engine for an aircraft may comprise:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox according to the ninth aspect, the gearbox configured        to receive an input from the core shaft and provide an output        drive to the fan so as to drive the fan at a lower rotational        speed than the core shaft.

Where the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

According to a tenth aspect there is provided a method of operating thegas turbine engine, the method comprising operating the aircraft gasturbine engine at maximum take-off conditions such that an eccentricityratio of each journal bearing, defined as 1-2H_(min)/c where H_(min) isa minimum oil film thickness between the internal and external slidingsurfaces and c is the diametral clearance of the journal bearing, may begreater than around 0.84.

According to an eleventh aspect there is provided a gearbox for anaircraft gas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and    -   a ring gear surrounding and engaged with the plurality of planet        gears, each of the plurality of planet gears being rotatably        mounted around a pin of a planet gear carrier with a journal        bearing having an internal sliding surface on the planet gear        and an external sliding surface on the pin, each journal bearing        comprising an oil flow path passing through the journal bearing        from an inlet to an outlet,    -   wherein, with the aircraft gas turbine engine operating at        maximum take-off conditions, a temperature of oil passing        through each oil flow path increases by between 15 and 30° C.        from the inlet to the outlet and a temperature of the oil at the        inlet is less than 105° C.

The ring gear may have a pitch circle diameter of around 550 mm orgreater. With the aircraft gas turbine engine operating at maximumtake-off conditions, the temperature of oil passing through each oilflow path may increase by between 15 and 25° C., or between 15 and 20°C., from the inlet to the outlet.

With the aircraft gas turbine engine operating at maximum take-offconditions, a specific oil flow rate through each oil flow path, definedas a flow rate of oil through the oil flow path divided by a diameterand length of the respective journal bearing may be less than around2000 l min⁻¹ m⁻².

In some examples the gearbox may be a planetary gearbox, i.e. where thering gear is connected to an output shaft and the sun gear connected toan input shaft.

With the aircraft gas turbine engine operating at maximum take-offconditions, a specific oil flow rate through each oil flow path, definedas a flow rate of oil through the oil flow path divided by a diameterand length of the respective journal bearing may be less than around1000 l min⁻¹ m⁻².

In some examples the gearbox may be a star gearbox, i.e. where an outputshaft is connected to a planet carrier connected to each planet gear andan input shaft is connected to the sun gear.

With the aircraft gas turbine engine operating at maximum take-offconditions, the specific oil flow rate through each oil flow path may begreater than around 400 l min⁻¹ m⁻².

With the aircraft gas turbine engine operating at maximum take-offconditions, a specific operating load multiplied by an operating slidingspeed of each journal bearing may be around 250 MPa m/s or greater.

With the aircraft gas turbine engine operating at maximum take-offconditions, the specific operating load multiplied by the operatingsliding speed of each journal bearing may be up to around 450 MPa m/s.

With the aircraft gas turbine engine operating at maximum take-offconditions, a specific operating load multiplied by an operating slidingspeed of each journal bearing may be around 450 MPa m/s or greater.

With the aircraft gas turbine engine operating at maximum take-offconditions, the specific operating load multiplied by the operatingsliding speed of each journal bearing may be up to around 720 MPa m/s.

A gas turbine engine for an aircraft may comprise:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox according to the eleventh aspect, the gearbox        configured to receive an input from the core shaft and provide        an output drive to the fan so as to drive the fan at a lower        rotational speed than the core shaft.

Where the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

According to a twelfth aspect there is provided a method of operatingthe gas turbine engine, the method comprising operating the aircraft gasturbine engine at maximum take-off conditions such that a temperature ofoil passing through each oil flow path increases by between 15 and 30°C. from the inlet to the outlet and a temperature of the oil at theinlet is less than 105° C.

Optional features according to the eleventh aspect may also apply to themethod of the twelfth aspect.

According to a thirteenth aspect there is provided a gearbox for anaircraft gas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and    -   a ring gear surrounding and engaged with the plurality of planet        gears, each of the plurality of planet gears being rotatably        mounted around a pin of a planet gear carrier with a journal        bearing having an internal sliding surface on the planet gear        and an external sliding surface on the pin,    -   wherein a diameter of each journal bearing divided by a pitch        circle diameter of the respective planet gear is less than        around 55%.

The ring gear may have a pitch circle diameter of around 550 mm orgreater.

The diameter of each journal bearing divided by the pitch circlediameter of the respective planet gear may be greater than around 50%.

With the aircraft gas turbine engine operating at maximum take-offconditions, a sliding speed of each journal bearing may be betweenaround 30 m/s and around 40 m/s.

With the aircraft gas turbine engine operating at maximum take-offconditions, a specific operating load multiplied by an operating slidingspeed of each journal bearing may be around 400 MPa m/s or greater.

With the aircraft gas turbine engine operating at maximum take-offconditions, the specific operating load multiplied by the operatingsliding speed of each journal bearing may be up to around 720 MPa m/s.

The internal or external sliding surface of the journal bearing may havea surface coating comprising a layer of an alloy having aluminium orcopper as a primary constituent.

A gas turbine engine for an aircraft may comprise:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox according to the thirteenth aspect, the gearbox        configured to receive an input from the core shaft and provide        an output drive to the fan so as to drive the fan at a lower        rotational speed than the core shaft.

Where the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

In accordance with a fourteenth aspect there is provided a method ofoperating the gas turbine engine, the method comprising operating theengine core to drive the core shaft and providing an output drive fromthe gearbox to the fan to drive the fan at a lower rotational speed thanthe core shaft.

Optional features relating to the thirteenth aspect may also apply tothe method of the fourteenth aspect.

According to a fifteenth aspect there is provided a gearbox for anaircraft gas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and    -   a ring gear surrounding and engaged with the plurality of planet        gears, each of the plurality of planet gears being rotatably        mounted around a pin of a planet gear carrier with a journal        bearing having an internal sliding surface on the planet gear        and an external sliding surface on the pin,    -   wherein, with the aircraft gas turbine engine operating at        maximum take-off conditions, a minimum oil film thickness        H_(min) between the internal and external sliding surfaces is a        function of a temperature T of oil flowing into the journal        bearing, such that H_(min)<B-AT, where A is 0.139 μm/° C. and B        is 27.8 μm.

The ring gear may have a pitch circle diameter of around 550 mm orgreater.

In some examples H_(min) may be greater than 2.3 μm.

In some examples, H_(min)>B-AT, where A is 0.034 μm/° C. and B is 6.4μm.

In some examples, H_(min)<B-AT, where A is 0.117 μm/° C. and B is 22 μm.

T may be greater than around 60° C., optionally greater than around 100°C. T may be less than around 120° C.

A gas turbine engine for an aircraft may comprise:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox according to the fifteenth aspect, the gearbox        configured to receive an input from the core shaft and provide        an output drive to the fan so as to drive the fan at a lower        rotational speed than the core shaft.

Where the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

According to a sixteenth aspect, there is provided a method of operatingthe gas turbine engine, the method comprising operating the aircraft gasturbine engine at maximum take-off conditions, a minimum oil filmthickness H_(min) between the internal and external sliding surfacesbeing a function of a temperature T of oil flowing into the journalbearing, such that H_(min)<B-AT, where A is 0.139 μm/° C. and B is 27.8μm.

Optional features according to the fifteenth aspect may also apply tothe method of the sixteenth aspect.

According to a seventeenth aspect there is provided a gearbox for anaircraft gas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and    -   a ring gear surrounding and engaged with the plurality of planet        gears, each of the plurality of planet gears being rotatably        mounted around a pin of a planet gear carrier with a journal        bearing having an internal sliding surface on the planet gear        and an external sliding surface on the pin,    -   wherein, with the aircraft gas turbine engine operating at        maximum take-off conditions, an eccentricity ratio, E, of each        journal bearing is a function of a temperature T of oil flowing        into the journal bearing, such that E>AT+B where A is        0.0015/° C. and B is 0.69.

The ring gear may have a pitch circle diameter of around 550 mm orgreater.

The eccentricity ratio may be defined as 1-2H_(min)/c, where H_(min) isa minimum oil film thickness between the internal and external slidingsurfaces and c is the diametral clearance of the journal bearing.

In some examples E may be less than around 0.98.

In some examples E<AT+B where A is 0.00033/° C. and B is 0.94.

In some examples E>AT+B where A is 0.00083/° C. and B is 0.84.

T may be greater than around 60° C., optionally greater than around 100°C.

T may be less than around 120° C.

A gas turbine engine for an aircraft may comprise:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox according to the seventeenth aspect, the gearbox        configured to receive an input from the core shaft and provide        an output drive to the fan so as to drive the fan at a lower        rotational speed than the core shaft.

Where the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

According to an eighteenth aspect there is provided a method ofoperating the gas turbine engine, the method comprising operating theaircraft gas turbine engine at maximum take-off conditions, aneccentricity ratio, E, of each journal bearing being a function of atemperature T of oil flowing into the journal bearing, such that E>AT+Bwhere A is 0.0015/° C. and B is 0.69.

Optional features according to the seventeenth aspect may also apply tothe method of the eighteenth aspect.

In accordance with a nineteenth aspect there is provided a gearbox foran aircraft gas turbine engine, the gearbox comprising:

-   -   a sun gear;    -   a plurality of planet gears surrounding and engaged with the sun        gear; and    -   a ring gear surrounding and engaged with the plurality of planet        gears, each of the plurality of planet gears being rotatably        mounted around a pin of a planet gear carrier with a journal        bearing having an internal sliding surface on the planet gear        and an external sliding surface on the pin,    -   wherein, with the aircraft gas turbine engine operating at        maximum take-off conditions, a Sommerfeld number of each journal        bearing is greater than around 4.

The ring gear may have a pitch circle diameter of around 550 mm orgreater.

An inefficiency of each journal bearing, defined as a percentage powerloss under maximum take-off conditions, may be less than around 0.225%.

A diametral clearance of each journal bearing may be between around 1‰and 2‰.

The diametral clearance of each journal bearing may be between around1.4‰ and 1.6‰.

With the aircraft gas turbine engine operating at maximum take-offconditions, a temperature of oil flowing into each journal bearing maybe less than or equal to around 100° C.

With the aircraft gas turbine engine operating at maximum take-offconditions, a pressure of oil flowing into each journal bearing atmaximum take-off conditions may be within a range from around 50 kPa toaround 350 kPa.

A gas turbine engine for an aircraft may comprise:

-   -   an engine core comprising a turbine, a compressor, and a core        shaft connecting the turbine to the compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades; and    -   a gearbox according to the nineteenth aspect, the gearbox        configured to receive an input from the core shaft and provide        an output drive to the fan so as to drive the fan at a lower        rotational speed than the core shaft.

Where the turbine is a first turbine, the compressor is a firstcompressor, and the core shaft is a first core shaft, the engine coremay further comprise a second turbine, a second compressor, and a secondcore shaft connecting the second turbine to the second compressor, thesecond turbine, second compressor, and second core shaft being arrangedto rotate at a higher rotational speed than the first core shaft.

According to a twentieth aspect there is provided a method of operatingthe gas turbine engine, the method comprising operating the aircraft gasturbine engine at maximum take-off conditions, wherein a Sommerfeldnumber of each journal bearing is greater than around 4.

The sliding speed of each journal bearing, according to any of the aboveaspects, may at maximum take-off conditions be 30, 31, 32, 33, 34, 35,36, 37, 38, 39, 40, 41, 42, 43, 44, 45, 46, 47, 48, 49 or 50 m/s orwithin a range defined by any two of the aforementioned values.

The specific operating load of each journal bearing, according to any ofthe above aspects, may at maximum take-off conditions be 5, 6, 7, 8, 9,10, 11, 12, 13, 14, 15, 16, 17, 18, 19, 20. 21. 22. 23. 24 or 25 MPa orwithin a range defined by any two of the aforementioned values.

The gas turbine engine may, in each of the above aspects, comprise aturbine, a compressor, a core shaft connecting the turbine to thecompressor and a fan located upstream of the engine core, the fancomprising a plurality of fan blades. The fan may have a moment ofinertia of between around 5.5×10⁷ and 9×10⁸ kg m², or alternativelybetween around 7.4×10⁷ and 7×10⁸ kg m², or alternatively between around8.3×10⁷ and 6.5×10⁸ kg m². The same features may also apply to the otheraspects of the invention.

The gearbox may, in each of the above aspects, have a gear ratio of 3.2to 4.5, and optionally 3.2 to 4.0. The gearbox may be in a starconfiguration.

The gas turbine engine may, in each of the above aspects, have aspecific thrust from 70 to 90 N kg⁻¹ and/or a bypass ratio at cruiseconditions of 12.5 to 18 or 13 to 16.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

The gearbox receives an input from the core shaft and outputs drive tothe fan so as to drive the fan at a lower rotational speed than the coreshaft. The input to the gearbox may be directly from the core shaft, orindirectly from the core shaft, for example via a spur shaft and/orgear. The core shaft may rigidly connect the turbine and the compressor,such that the turbine and compressor rotate at the same speed (with thefan rotating at a lower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches),260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm(around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160inches) or 420 cm (around 165 inches). The fan diameter may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allunits in this paragraph being J kg⁻¹K⁻¹/(ms⁻¹)²). The fan tip loadingmay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 0.28 to 0.31, or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypassduct may be substantially annular. The bypass duct may be radiallyoutside the core engine. The radially outer surface of the bypass ductmay be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110 Nkg⁻¹ s, 105 N kg⁻¹ s, 100 N kg⁻¹ s, 95 N kg⁻¹s, 90 N kg⁻¹ s, 85 N kg⁻¹ sor 80 N kg⁻¹s. The specific thrust may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 80 Nkg⁻¹sto 100 N kg⁻¹s, or 85 N kg⁻¹s to 95 N kg⁻¹s. Such engines may beparticularly efficient in comparison with conventional gas turbineengines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmay be formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions may mean cruise conditions of anaircraft to which the gas turbine engine is attached. Such cruiseconditions may be conventionally defined as the conditions atmid-cruise, for example the conditions experienced by the aircraftand/or engine at the midpoint (in terms of time and/or distance) betweentop of climb and start of decent.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be the cruise condition. For someaircraft, the cruise conditions may be outside these ranges, for examplebelow Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions at an altitude that is in the range offrom 10000 m to 15000 m, for example in the range of from 10000 m to12000 m, for example in the range of from 10400 m to 11600 m (around38000 ft), for example in the range of from 10500 m to 11500 m, forexample in the range of from 10600 m to 11400 m, for example in therange of from 10700 m (around 35000 ft) to 11300 m, for example in therange of from 10800 m to 11200 m, for example in the range of from 10900m to 11100 m, for example on the order of 11000 m. The cruise conditionsmay correspond to standard atmospheric conditions at any given altitudein these ranges.

Purely by way of example, the cruise conditions may correspond to: aforward Mach number of 0.8; a pressure of 23000 Pa; and a temperature of−55 degrees C. Purely by way of further example, the cruise conditionsmay correspond to: a forward Mach number of 0.85; a pressure of 24000Pa; and a temperature of −54 degrees C. (which may be standardatmospheric conditions at 35000 ft).

As used anywhere herein, “cruise” or “cruise conditions” may mean theaerodynamic design point. Such an aerodynamic design point (or ADP) maycorrespond to the conditions (comprising, for example, one or more ofthe Mach Number, environmental conditions and thrust requirement) forwhich the fan is designed to operate. This may mean, for example, theconditions at which the fan (or gas turbine engine) is designed to haveoptimum efficiency.

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

As used herein, a maximum take-off (MTO) condition has the conventionalmeaning. Maximum take-off conditions may be defined as operating theengine at International Standard Atmosphere (ISA) sea level pressure andtemperature conditions+15° C. at maximum take-off thrust at end ofrunway, which is typically defined at an aircraft speed of around 0.25Mn, or between around 0.24 and 0.27 Mn. Maximum take-off conditions forthe engine may therefore be defined as operating the engine at a maximumtake-off thrust (for example maximum throttle) for the engine at ISA sealevel pressure and temperature+15° C. with a fan inlet velocity of 0.25Mn.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is a schematic transverse sectional view across an example planetgear mounted on a pin of a planet gear carrier with a journal bearing;

FIG. 5 is a schematic longitudinal sectional view through the exampleplanet gear of FIG. 4;

FIG. 6 is a partial sectional view through an external surface of anexample planet gear journal bearing;

FIG. 7 is a schematic drawing of a transverse section of an exampleplanet gear, showing an exaggerated oil film thickness variation;

FIG. 8 is a schematic plot of operating specific load as a function ofsliding speed for a number of example gearboxes;

FIG. 9 is a schematic plot of eccentricity ratio as a function ofpercentage loading for a number of example gearboxes;

FIG. 10 is a schematic plot of inefficiency as a function ofeccentricity ratio of journal bearings for a range of example gearboxesoperating at maximum take-off conditions;

FIG. 11 is a schematic plot of minimum oil film thickness as a functionof oil inlet temperature for a range of example gearboxes operating atmaximum take-off conditions;

FIG. 12 is a schematic plot of eccentricity ratio as a function of oilinlet temperature for a range of example gearboxes operating at maximumtake-off conditions;

FIG. 13 is a schematic plot of inefficiency as a function of Sommerfeldnumber for a range of example gearboxes operating at maximum take-offconditions; and

FIG. 14 is a schematic plot of specific oil flow as a function of PV fora range of example gearboxes operating at maximum take-off conditions.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually orthogonal.

FIG. 4 illustrates schematically an example planet gear 32 of the typeshown in the example epicyclic gearbox 30 of FIG. 3. The planet gear 32is mounted around a pin 41, the inner surface of the planet gear 32 andthe outer surface of the pin 41 forming sliding surfaces of a journalbearing 42, allowing the planet gear 32 to rotate relative to the pin41. In use, the sliding surfaces are lubricated with oil to allow theplanet gear 32 and pin 41 to rotate smoothly relative to each other. Thegap between the sliding surfaces is shown exaggerated in FIG. 4 forclarity.

The sliding surfaces of the journal bearing 42 in the example of FIG. 4are shown as the inner surface of the planet gear 32 and outer surfaceof the pin 41. In alternative examples, a sleeve may be provided aroundthe pin 41, an outer surface of which provides the inner sliding surfaceof the journal bearing 42, the sleeve being fixed to the outer surfaceof the pin 41, for example by an interference fit. A bush (or bushing)may alternatively or additionally be provided, the inner surface ofwhich provides the outer sliding surface of the journal bearing 42, thebush being fixed to the inner surface of the planet gear 32, for exampleby an interference fit between the bush and the gear 32. An advantage offorming the sliding surfaces of the journal bearing 42 from the pin 41and planet gear 32 themselves is that tolerances of the journal bearing42 can be more tightly controlled, while an advantage of using one orboth of a sleeve and a bush is that the journal bearing may be morereadily repaired by replacing one or both components when worn.

The planet gear 32 is defined by an inner surface diameter 43, which mayalso be defined as the diameter of the journal bearing 42, and an outerpitch circle diameter 44. The planet gear 32 comprises a plurality ofteeth 45 extending around the outer circumference of the gear 32. Thetotal number of teeth 45 may differ from that shown in FIG. 4 dependingon the specifics of the application. The teeth 45 may be arranged in aspur gear or helical gear form, i.e. either parallel or at an angle tothe rotational axis of the gear 32. A helical gear form is a more commonarrangement because this allows for a smoother transition between thegear teeth 45 of the planet gear 32 and corresponding teeth on the ringgear 38 and sun gear 28 (FIG. 3) as the gears rotate relative to eachother.

FIG. 5 illustrates schematically an axial section through the planetgear 32 and pin 41 of FIG. 4. The transverse section shown in FIG. 4 istaken along the line A-A′ indicated in FIG. 5. The pin 41 is mounted tothe planet carrier 34, in this example by extending through thethickness of the planet carrier 34. The pin 41 may be fixed to thecarrier 34 by welding, bolting or by otherwise securing the pin 41 andcarrier 34 to prevent relative movement between the pin 41 and carrier34 when in use. In operation, forces are transmitted between the pin 41and carrier 34 primarily through shear forces on the pin 41 transverseto the axis 51 of the pin 41, which also result in bending momentsapplied to the pin 41 along the axis 51. In a star gearbox arrangement,in which the planet carrier 34 is fixed relative to the engine frame,the net forces on the planet gears 32 act in a direction tangential to adiameter of the planet gear 32 centres. In a planetary gearboxarrangement, in which the outer ring gear 38 (FIG. 3) is fixed, the netforces on the planet gears 32 are tilted towards the centre of the sungear due to the additional centripetal force component required tomaintain the planet gears 32 rotating about the sun gear 28, thecentripetal force being a function of the rotational speed of the planetcarrier 34. An advantage of the gearbox being configured in a stararrangement is that loading on the pins is reduced when the gearbox isoperating at high speeds.

The planet gear 32 is shown in FIG. 5 with a journal bearing portion 52having a length L smaller than a total width 54 of the planet gear 32.The length L of the journal bearing 42 may be selected according to theloads experienced during operation of the gearbox and to optimise aratio between the journal bearing length L and the diameter D of thejournal bearing 42. The diameter D may be defined by either the outersliding surface, corresponding to the inner surface of the planet gear32 in the example shown in FIG. 5, by the inner sliding surface,corresponding to the outer surface of the pin 41, or by a mean diameterbetween the two. In practice, the difference between the two diameters,termed the diametral clearance c (the distance c/2 being shown in FIG.4), is small, typically within less than 0.5% of either diameter. For anexample range of diameters of between 120 mm and 200 mm, the differencemay be between around 0.1% and 0.3%, i.e. between around 120 μm andaround 600 μm, with a typical diametral clearance of around 150 μm.

The length 52 of the journal bearing 42 may in some examples be the sameas, or greater than, the total width of the planet gear 32.

In particular examples, a ratio L/D of the length L of the journalbearing 42 to the diameter D of the journal bearing 42 may be in a rangefrom around 0.5 to 1.4, optionally between around 1.1 and 1.3. A lowerL/D ratio reduces misalignment of the gears 32 relating to the pins 41,in part by reducing the bending moment applied to the pins, therebykeeping the pins 41 more parallel with the gears 32. The L/D ratioshould, however, be kept above around 0.5, or optionally around 1.1, toavoid the specific loading on the journal bearing from becoming too highand adversely affecting the lifetime of the bearing.

FIG. 6 illustrates schematically an example structure of a surfacecoating 61 that may be applied to either sliding surface of the journalbearing 42. The underlying material 62 may be either the pin 41 or thering gear 32, or in alternative examples may be a sleeve or bush of thetype described above. The overall thickness of the surface coating 61may be in the region of between 40 and 200 micrometres thick, with aspecific example thickness in the region of around 100 micrometres.

Although the surface coating 61 may be applied to either surface of thejournal bearing 42, applying the coating 61 to the outer surface of thepin 41 may in practice be preferable due to practical limitations ofdeposition methods for internal surfaces. Common deposition methods suchas physical vapour deposition (PVD) may be more suitable for applicationof coatings to an external rather than internal surface. Othertechniques such as casting may be more applicable for application of acoating to an internal surface, although casting is generally lesssuitable for creating a coating of the thickness range defined above,and with the tolerances required for journal bearings.

An example surface coating 61 may comprise three layers 61 a-c. A firstlayer 61 a is deposited that has a thermal expansion coefficient betweenthat of the underlying material 62 and the second layer 61 b. With steelas the underlying material, the first layer 61 a may for example be acopper-based alloy. The second layer 61 b, which typically forms thelargest thickness layer in the surface coating 61, i.e. having athickness of between around 50% and 95% of the total thickness of thesurface coating 61, may be composed of a copper- or aluminium-basedalloy, i.e. a metallic alloy having either copper or aluminium as aprimary constituent, an example being a leaded bronze, i.e. an alloy ofcopper, lead and tin. Such an alloy is selected to have a lower hardnesscompared with that of the material forming the other surface of thejournal bearing, so that any particles that are not filtered out fromthe oil may instead become embedded in the second layer 61 b, reducingtheir ability to wear the surfaces of the journal bearing.

The third layer 61 c may be one that is considerably thinner than thefirst and second layers 61 a, 61 b and composed of a material having alower hardness than the second layer 61 b, for example a lead-basedalloy. The third layer acts to reduce friction between the surfaces ofthe journal bearing, particularly when starting from a stationaryposition where an oil layer between the surfaces has not been built up.The third layer 61 c may for example have a thickness of between 1 and10 micrometres.

In a particular example, the first layer 61 a may be between 10 and 20micrometres in thickness, the second layer between around 40 and 100micrometres in thickness and the third layer between around 1 and 15micrometres or between 1 and 10 micrometres or between 5 and 15micrometres or between 10 and 15 micrometres in thickness.

The bearing materials have been developed to provide an optimumcompromise between ‘hard/strong’ and ‘soft/flexible’ for the specificapplication in a gearbox for a gas turbine engine. ‘Hard’ propertiesaddress the requirements of contact wear resistance, fatigue, and loadcarrying capacity. ‘Soft’ properties are advantageous to providecompatibility (to the countersurface), conformability, and embeddabilityof the surface. It has been found that this may help to ensure continuedoperation in imperfect conditions. In addition, the proposed arrangementhas been developed to address environmental factors such as corrosionand oxidation resistance

In a specific example, the coating layer 61 b may be an(aluminium-tin-copper alloy (for example SAE783). In another specificexample, a leaded bronze alloy (such as SAE49) may be used for thecoating layer 61 b. Such alloys have been found to be suitable for anarduous duty cycle, for example where the maximum operating specificload multiplied by the maximum operating sliding speed is around 240 MPam/s or greater. The soft properties at the running surface can befurther enhanced with a thin overlay coating 61 c (for example up toaround 12 μm) such as SAE 194 lead-indium without compromising the loadcarrying capacity of the underlying material.

FIG. 7 illustrates a schematic cross-sectional view of a gear 32 mountedaround a pin 41, forming a journal bearing 42 between the outer surfaceof the pin 41 and the inner surface of the gear 32. A difference ininner and outer diameter between the gear 32 and pin 41 respectively,i.e. the diametral clearance, is exaggerated to show a variation in oilfilm thickness that arises when the gearbox is in operation. The oilfilm thickness is lower over a portion 71 of the journal bearing where aload F is transferred between the pin 41 and the gear 32, for example inthe direction indicated by arrow 72. A specific loading on the journalbearing 42 is defined by the load F on the bearing 42 applied over anarea defined by the length L and diameter D of the bearing 42, i.e. thespecific loading is F/LD, typically measured in MPa or N/mm².

An oil flow path through the journal bearing 42 passes through a centralbore 73 of the pin 41 through an inlet passage 74 and into a clearancebetween the pin 41 and gear 32. The oil flows around the journalbearing, dragged through the minimum clearance by the relative rotationbetween the pin 41 and gear 32, and exits via the edges of the bearing42. Oil is cooled and recirculated via a scavenge and pump (not shown).The oil flowing into the journal bearing may be pressurised to betweenaround 50 and 350 kPa (0.5 to 35 bar). A minimum oil pressure isrequired to provide sufficient oil to the bearing so that the area overwhich force is applied is covered with a supply of oil. Higher pressureswill tend to force greater amounts of oil through the bearing, but havediminishing effects on lubricating and cooling the bearing as greateramounts will tend to travel via the wider portion of the clearancebetween the pin 41 and gear 32 rather than via the minimum clearanceportion 71. Higher oil pressures will tend to reduce the temperaturedifference between the inlet and outlet oil flows, which makesextraction of heat more difficult, requiring larger heatsinks. Anoptimum oil flow pressure and temperature difference will therefore tendto be required to minimise on weight in relation to oil pumps andheatsinks. The pressure and temperature differences defined herein havethus been chosen to provide the required lubrication, but with asufficiently high temperature difference to enable sufficiently lowweight of heat exchangers to remove the heat. The low weight of heatexchanger may be a particularly important consideration for gearboxes tobe used in a gas turbine for an aircraft, because of the importance ofweight on the overall fuel consumption of the aircraft to which theengine is provided.

The dimensional and positional accuracy of the pins 41 and gears 32 ofthe gearbox will affect how the oil film thickness varies, as well asthe viscosity and temperature of the oil. To maintain a uniform oiltemperature across each journal bearing, symmetric oil feed paths may beprovided in the gearbox, and a plenum for mixing oil prior to being fedinto the gearbox may be sufficiently large to allow for a uniformtemperature of oil being fed into the gearbox at different feed points.As a result, a temperature variation between oil fed to each of thejournal bearings may be no more than 1 degree Celsius, for example withthe engine operating at cruise conditions. A variation in oil pressureis preferably also uniform between the journal bearings, but this willtypically have less effect than a variation in temperature because anincrease in pressure above a minimum required will tend to simply causemore oil to flow through the bearing, having minimal effect onoperation.

The operational oil film thickness, i.e. the thickness of the oil filmin each journal bearing during operation of the engine, may be definedas a proportion of the journal bearing diameter. The minimum operationaloil film thickness for each journal bearing during operation, forexample at MTO conditions, at which loading of the gearbox is at itshighest, may be less than around 8 micrometres for a journal bearingdiameter of between around 120 mm and 200 mm, and optionally greaterthan around 3.5 micrometres. The clearance of the journal bearing maytypically be between around 1 and 3‰ (0.1% and 0.3%) of the journalbearing diameter, for example around 1.5‰ (0.15%). The journal bearingdiameter may, as described above, be defined as the diameter of theinner sliding surface of the planet gear. A variation between theminimum operational thickness of each journal bearing, also for exampleat MTO conditions, may be less than around 8% of a mean minimum oil filmthickness. For example, if the mean minimum oil film thickness is around6 micrometres, the maximum difference between the minimum oil filmthickness across all of the journal bearings will be around +/−around0.5 micrometres.

The operational oil film thickness will, as illustrated schematically inFIG. 7, vary around each journal bearing between a minimum thickness ata point of maximum loading to a maximum thickness at a pointdiametrically opposite from the point of maximum loading. The point ofmaximum loading will tend to follow a linear path along the length ofthe journal bearing parallel to its axis of rotation. A ratio betweenthe maximum and minimum oil film thickness will be highest duringmaximum take-off conditions and will be around 1 at idle conditions,i.e. with no significant load being transferred across the gearbox.

FIG. 8 illustrates an example plot of operating specific load (y axis,in MPa) as a function of sliding speed (x axis, in m/s) for journalbearings in a range of example gearboxes of the type disclosed herein,each operating under maximum take-off conditions. The specific loadingfor a journal bearing is as defined above. The sliding speed for ajournal bearing is defined as the relative tangential speed of the innerand outer surfaces of the journal bearing. The dotted lines 81 a, 81 b,81 c represent constant values for a multiple of operating specificloading and sliding speed, which may be termed PV (being a multiple ofpressure and velocity), of 200, 400 and 600 MPa m/s respectively.

At higher specific loads or sliding speeds, or higher values of PV ingeneral, a surface coating comprising a layer of an alloy havingaluminium or copper as a primary constituent, for example forming thesecond layer 61 b as shown in FIG. 6, may be used. Copper as a primaryconstituent may be preferable for higher diameter journal bearings, forexample greater than 120 mm in diameter.

The maximum operating specific loading of each journal bearing in thegearbox may be greater than 5 MPa, or may be greater than any one of 6MPa, 7 MPa, 8 MPa, 9 MPa, 10 MPa, 11 MPa, 12 MPa, 13 MPa, 14 MPa, 15MPa, 16 MPa or 17 MPa. The maximum sliding speed of the journal bearingsmay be defined by the corresponding sliding speed for the curves 81 a-cshown in FIG. 8. In particular examples, the maximum operating specificloading may be around 13 MPa or around 18 MPa, with a sliding speed ineach case of around 42 or 38 m/s, indicated as data points 82, 83respectively on FIG. 8. In a general aspect, the specific loading may bewithin a range from around 10 to 20 MPa and the sliding speed within arange from around 35 to 45 m/s at maximum take-off conditions. Theseranges may apply in particular for a planetary gearbox arrangement.

Points 82, 83 represent specific pressure and sliding speed values atmaximum take-off conditions for journal bearings in two exampleplanetary gearboxes, with journal bearing diameters of around 155 and140 mm respectively and journal bearing L/D ratios of around 1.11 and1.24 respectively, both with a diametral clearance of around 1.5‰. ThePV values at maximum take-off conditions for points 82 and 83 are around560 and 650 MPa m/s respectively.

Points 84, 85 and 86 in FIG. 8 represent specific pressure and slidingspeed values at maximum take-off conditions for journal bearings inthree example star gearboxes of different sizes, with journal bearingdiameters of around 100, 120 and 180 mm respectively and journal bearingL/D ratios of around 1.45, 1.35 and 1.13 respectively, each with adiametral clearance of around 1.5‰. The PV values at maximum take-offconditions for points 84, 85 and 86 are around 325, 335 and 370 MPa m/s.In a general aspect, the specific loading for such examples may bewithin a range from around 5 to 10 MPa and the sliding speed within arange from around 45 to 55 m/s at maximum take-off conditions. The PVvalues may be in a range having a lower limit of any one of 200, 220,240, 260, 230 or 300 MPa m/s and an upper limit of any one of 310, 330,350, 370, 390, 410 or 430 MPa m/s.

In a further general aspect therefore, the specific loading for theabove-mentioned examples may be within an overall range from around 5 to20 MPa and the sliding speed within a range from around 30 or 35 to 50or 55 m/s at maximum take-off conditions.

The higher specific loads for the planetary gearbox journal bearings(points 82, 83) partly reflect the additional centripetal loading oneach journal bearing due to the rotation of each planet gear about thecentral sun gear, while the planet gears in the star gearboxes (points84, 85, 86) do not rotate about the central sun gear.

The y-axis spread of specific load on each of the data points 82-86represents the variation in specific load over a +/−10% variation intorque load around a nominal torque load at maximum take-off conditions.

An upper limit for PV may be around 720 MPa m/s, while a lower limit maybe around 240 or 300 MPa m/s. Upper limits may alternatively be definedby an upper limit for one or both of the sliding speed and operatingspecific load, for example an upper limit of around 45, 50, 55 or 60 m/sfor the sliding speed and an upper limit of around 10, 20 or 30 MPa forthe operating specific loading. Lower limits may be defined by slidingspeeds of around 30, 35, 40 or 45 m/s, or by specific loads of around 5or 10 MPa, among others specified herein.

The eccentricity ratio of a journal bearing during operation of the gasturbine engine, for example while operating at MTO conditions, isdefined as 1-2H_(min)/c, where H_(min) is the minimum oil film thickness(shown in FIG. 7) and c the diametral clearance (shown in FIG. 4, withthe gear 32 and pin 41 arranged concentrically). FIG. 9 illustrates thevariation in eccentricity ratio (y axis) for a range of examplegearboxes as a function of percentage of the journal bearing design load(x axis). First, second and third example star gearboxes 91,92, 93(corresponding to the same gearbox designs having data points 84, 85, 86respectively in FIG. 8) exhibit a variation in eccentricity ratio ofbetween around 0.2 and 0.3 between 90% and 110% of design load, and haveeccentricity ratios that range between around 0.79 and 0.91 over thisrange of design loads. Eccentricity ratios 94, 95 for first and secondexample planetary gearboxes (corresponding to gearbox designs havingdata points 82 and 83 in FIG. 8), having higher absolute design loads,are between around 0.94 and 0.97 over a similar design load range, withthe eccentricity ratio varying within this range by between around 0.03and around 0.05.

The diametral clearance, c, may be within a range of between around 1and 2‰, i.e. between around 0.1 and 0.2%. A smaller diametral clearancewill tend to increase the area over which the pressure between the innerand outer surfaces of the journal bearing is distributed, but this willbe in combination with a narrower path through which the oil through thebearing is forced as the bearing rotates, limiting the flow rate of oilthrough the bearing and ultimately causing the bearing to seize as thediametral clearance is reduced further. A higher diametral clearancewill tend to reduce the area over which the pressure is distributed butwill also make travel of the oil through the bearing easier. An optimumbalance between the factors is therefore required which, particularlyfor higher eccentricity ratios of between around 0.94 and 0.97, may bebetween around 1 and 2‰, and optionally between around 1.4 and 1.6 ‰.

FIG. 10 is a schematic plot of inefficiency as a function ofeccentricity ratio for the above-mentioned range of gearboxes. Theexample star gearboxes tend to have higher inefficiencies, rangingbetween around 0.17 and 0.3%, while the example planetary gearboxes haveinefficiencies between around 0.1 and 0.16%. The eccentricity ratiosrange between values as stated above. The variation of inefficiencyversus eccentricity follows the general trend 1001 shown in FIG. 10,with a higher eccentricity ratio resulting in a higher efficiency, i.e.a lower inefficiency. A range of eccentricity ratios may be aspreviously stated, while a range of inefficiency may be less than around0.225%, and may be between around 0.225 and around 0.1%. Increasing theeccentricity ratio further will tend to increase the risk of the journalbearing seizing due to the minimum thickness of the oil film becomingtoo small to sustain an oil film separating the pin and gear under therequired range of loading.

FIG. 11 illustrates a series of trendlines of H_(min) (in μm) as afunction of oil inlet temperature (in ° C.) for the above-mentionedexample star and planetary gearboxes, all operating at maximum take-offconditions. The star gearboxes (lines 1101, 1102, 1103) tend to havehigher H_(min) values over the range of temperatures and with trendlineshaving a steeper gradient, while the planetary gearboxes (lines 1104,1105) tend to have lower H_(min) values and with more shallow gradients.Each trendline tends to follow a function of the form H_(min)=B-AT,where T is temperature (in ° C.) and A and B are constants that arecharacteristic of the particular gearbox design. Except for one of thestar gearboxes, the minimum oil film thickness H_(min) at maximumtake-off conditions for each gearbox is within a region having an upperbound defined by the line 1106, where A is 0.139 μm/° C. and B is 27.8μm. A minimum value of H_(min) may be around 2.3 μm, below which the oilfilm may be insufficient to prevent seizing of the journal bearing. Twofurther lines 1107, 1108 define further upper and lower boundsrespectively, with line 1107 defined by A=0.117 μm/° C. and B=22 μm andline 1108 defined by A=0.034 μm/° C. and B=6.4 μm. An overall range forthe inlet oil temperature may be between 60 and 120° C., with anoptional range of greater than around 100° C. and less than around 120°C. At lower temperatures the oil viscosity increases, reducinglubrication efficiency, whereas at higher temperatures the resultinglower oil viscosity may cause the minimum oil film thickness to becometoo small.

FIG. 12 illustrates a series of trendlines of eccentricity ratio, E, ofjournal bearings of the various example gearboxes as a function of oilinlet temperature, with the gearbox in each case operating at maximumtake-off conditions. The star gearboxes (lines 1201, 1202, 1203) tend tohave lower values of E over the entire temperature range and trendlineshaving steeper gradients, while the planetary gearboxes (lines 1204,1205) tend to have higher values of E and more shallow gradients. Ineach case the trendline tends to follow a function of the form E=AT+Bwhere T is the oil inlet temperature and A and B constants. A maximumvalue for E may be around 0.98, above which the oil film may be toosmall to sustain lubrication of the journal bearing. The eccentricityratio E may be above a trendline 1206 defined by A=0.0015/° C. andB=0.69, or alternatively may be above a trendline 1207 defined byA=0.00083/° C. and B=0.84, and may be below a trendline 1208 defined byA=0.00033/° C. and B=0.94. As for the examples in FIG. 11, an overallrange for the inlet oil temperature may be between 60 and 120° C., withan optional range of greater than around 100° C. and less than around120° C.

The Sommerfeld number, S, of a journal bearing is defined as:

$S = {( \frac{d}{c} )^{2}\frac{\mu N}{P}}$where d is the outer diameter of the pin 41 (FIG. 7), c is the diametralclearance, p is the absolute viscosity of the lubricant, N the relativerotational speed of the journal bearing (in revolutions per second) andP is the loading applied across the projected bearing area, i.e. F/LD,where L is the journal bearing length and D the diameter.

A higher PV value, resulting in a lower inefficiency value, will tend toincrease the Sommerfeld number for a journal bearing. FIG. 13illustrates a general relationship, given by trendline 1301, betweeninefficiency and Sommerfeld number for the above-mentioned range ofgearboxes. The example star type gearboxes tend to have journal bearingswith a lower Sommerfeld number, ranging between around 1 and 9 atmaximum take-off conditions. The two planetary gearboxes, with higher PVvalues, higher eccentricity and lower inefficiencies, have journalbearings that tend to have higher Sommerfeld numbers, ranging in generalbetween around 4 and 21 under varying oil temperature and loadings.Under more optimal conditions of oil temperature, the Sommerfeld numberfor these planetary gearbox journal bearings tends to be between around10 and 16. In a general aspect, the Sommerfeld number of each journalbearing may be greater than around 4, with an inefficiency of around0.225% or less under maximum take-off conditions. As mentioned above,the minimum inefficiency may be around 0.1%. The maximum Sommerfeldnumber may be around 21.

FIG. 14 shows a schematic relationship between specific oil flow, i.e.oil flow (in litres/minute) divided by an area of the journal bearing(i.e. LD, in m²) as a function of PV (in MPa m/s). The region 1401defined between upper and lower bounds 1402,1403 encompasses the journalbearing for each of the five above mentioned gearboxes, with the threestar gearboxes at the left hand end of the region 1401, generally belowaround 450 MPa m/s and above 240 MPa m/s and the two planetary gearboxesat the right hand end, generally in a region between around 450 and 720MPa m/s. The general relationship illustrates that a higher value for PVis associated with a lower specific oil flow, therefore requiring lowerpressure oil flows and a general reduction in weight of associatedequipment such as oil pumps and heatsinks for a given power rating. Ahigher PV value is therefore particularly advantageous for a gas turbineengine for an aircraft.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

The invention claimed is:
 1. A method of operating a gearbox for anaircraft gas turbine engine, the gearbox comprising: a sun gear; aplurality of planet gears surrounding and engaged with the sun gear; anda ring gear surrounding and engaged with the plurality of planet gears,the ring gear having a pitch circle diameter of around 550 mm orgreater, wherein each of the plurality of planet gears is rotatablymounted around a pin of a planet gear carrier with a journal bearinghaving an internal sliding surface on the planet gear and an externalsliding surface on the pin, an oil film between the internal surface onthe planet gear and the external sliding surface on the pin varyingbetween a maximum thickness and a minimum thickness around the journalbearing, the method comprising operating the aircraft gas turbine engineat maximum take-off conditions such that the minimum thickness of theoil film varies between the plurality of planet gears by no more than 8%from a mean minimum oil film thickness of the plurality of planet gears.2. The method of claim 1, wherein a diameter, D, of each journal bearingis between around 120 mm and around 200 mm.
 3. The method of claim 2,wherein a length, L, of the internal and external sliding surfaces ofeach journal bearing is between around 0.5 and around 1.4 of thediameter, D.
 4. The method of claim 3, wherein L/D is between around 1.1and around 1.3.
 5. The method of claim 1 wherein the mean minimum oilfilm thickness at maximum take-off conditions is between around 3.5 and8 micrometres.
 6. The method of claim 1, wherein an eccentricity ratioof each journal bearing during operation of the gas turbine engine atmaximum take-off conditions is within a range of between around 0.94 and0.97.
 7. The method of claim 1, wherein for each journal bearing in thegearbox a specific loading multiplied by a sliding speed is greater thanaround 240 MPa m/s.
 8. The method of claim 1, wherein the gas turbineengine comprises a turbine, a compressor, a core shaft connecting theturbine to the compressor and a fan located upstream of the engine core,the fan comprising a plurality of fan blades, the fan having a moment ofinertia of between around 5.5×10⁷ and 9×10⁸ kg m².
 9. A gearbox for anaircraft gas turbine engine, the gearbox comprising: a sun gear; aplurality of planet gears surrounding and engaged with the sun gear; anda ring gear surrounding and engaged with the plurality of planet gears,the ring gear having a pitch circle diameter of around 550 mm orgreater, each of the plurality of planet gears being rotatably mountedaround a pin of a planet gear carrier with a journal bearing having aninternal sliding surface on the planet gear and an external slidingsurface on the pin, an oil film between the internal surface on theplanet gear and the external sliding surface on the pin varying betweena maximum thickness and a minimum thickness around the journal bearing,and wherein, during operation of the aircraft gas turbine engine atmaximum take-off conditions, the minimum thickness of the oil filmvaries between the plurality of planet gears by no more than 8% from amean minimum oil film thickness.
 10. The gearbox of claim 9, wherein adiameter, D, of each journal bearing is between around 120 mm and around200 mm.
 11. The gearbox of claim 10, wherein a length, L, of theinternal and external sliding surfaces of each journal bearing isbetween around 0.5 and around 1.4 of the diameter.
 12. The gearbox ofclaim 9, wherein, during operation of the aircraft gas turbine engine atmaximum take-off conditions, the mean minimum oil film thickness isbetween around 3.5 and 8 micrometres.
 13. The gearbox of claim 9,wherein, during operation of the aircraft gas turbine engine at maximumtake-off conditions, an eccentricity ratio of each journal bearing iswithin a range of between around 0.94 and 0.97.
 14. The gearbox of claim9, wherein, during operation of the aircraft gas turbine engine atmaximum take-off conditions, for each journal bearing in the gearbox aspecific loading multiplied by a sliding speed is greater than around240 MPa m/s.
 15. The gearbox of claim 9, wherein the gearbox has a gearratio of 3.2 to 4.5 or 3.2 to 4.0.
 16. The gearbox of claim 9, whereinthe gearbox is in a star configuration.
 17. A gas turbine engine for anaircraft, comprising: an engine core comprising a turbine, a compressor,and a core shaft connecting the turbine to the compressor; a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades; and the gear box according to claim 9, the gearbox configured toreceive an input from the core shaft and provide an output drive to thefan so as to drive the fan at a lower rotational speed than the coreshaft.
 18. The gas turbine engine of claim 17, wherein: the turbine is afirst turbine, the compressor is a first compressor, and the core shaftis a first core shaft; the engine core further comprises a secondturbine, a second compressor, and a second core shaft connecting thesecond turbine to the second compressor; and the second turbine, secondcompressor, and second core shaft are arranged to rotate at a higherrotational speed than the first core shaft.
 19. The gas turbine engineaccording to claim 17, wherein the gas turbine engine has: a specificthrust from 70 to 90 N kg-1; and/or a bypass ratio at cruise conditionsof 12.5 to
 18. 20. The gas turbine engine according to claim 17,wherein: the fan has a moment of inertia of between around 5.5×10⁷ and9×10⁸ kg m².